PERFORMANCE ANALYSIS ON A TURBO JET
ENGINE
KISHORE S RAHUL KUMAR
Thermal and Propulsion Thermal and Propulsion
Indian Institute of Space Science and Technology Indian Institute of Space Science and Technology
Thiruvananthapuram, India Thiruvananthapuram, India
SC22M023 SC22M027
RAJATH ANKIT KUMAR SINHA
Thermal and Propulsion Thermal and Propulsion
Indian Institute of Space Science and Technology Indian Institute of Space Science and Technology
Thiruvananthapuram, India Thiruvananthapuram, India
SC22M028 SC22M020
MITHUNA LINGARAJ SAMALLA PRANITHA CHAKRIN
Thermal and Propulsion Thermal and Propulsion
Indian Institute of Space Science and Technology Indian Institute of Space Science and Technology
Thiruvananthapuram, India Thiruvananthapuram, India
SC22M024 SC22M029
SAJID MOMIN
Thermal and Propulsion
Indian Institute of Space Science and Technology
Thiruvananthapuram, India
SC22M019
Abstract—The experiment for analysing the performance of
SR30 turbojet is carried out in Thermal and Propulsion Lab
(IIST).The setup is coupled with data acquisition system to
store and view the data. Various parameters were analysed for
measuring the performance of the turbojet engine. Experiment
was carried out by varying the rpm by throttle regulator and
a fixed compressed air was supplied through compressor at
a pressure of 10 bar. The parameters acquired by the data
acquisition system during the experiment were thrust generated,
compressor and turbine inlet and exhaust temperature. Thrust
and TSFC at varying speeds were studied by plotting the graph
to analyse the performance of the engine test setup. Fig. 1: Basic components of a turbojet engine
Index Terms—
T SF C - Thrust specific fuel consumption
Dc - Diameter of the engine intake bell section
d - Diameter of the tacho generator housing nozzle. The gas turbine has an air inlet which includes inlet
a - Cross sectional area at pitot static tube location guide vanes, a compressor, a combustion chamber, and a
Tc - Temperature at Nozzle exit turbine (that drives the compressor). The compressed air from
Vc - Velocity of air at compressor inlet the compressor is heated by burning fuel in the combustion
M - Mach number of the airflow at this section chamber and then allowed to expand through the turbine.
F - Net Thrust of the engine
The turbine exhaust is then expanded in the propelling nozzle
where it is accelerated to high speed to provide thrust [1]. Two
engineers, Frank Whittle in the United Kingdom and Hans von
I. I NTRODUCTION Ohain in Germany developed the concept independently into
The turbojet is an airbreathing jet engine that is typically practical engines during the late 1930s. The turbojet engine
used in aircraft. It consists of a gas turbine with a propelling contains a diffuser at the front end and an expansion nozzle
Fig. 2: T-S diagram for ideal Turbojet Cycle Fig. 3: Photograph of the experimental set-up
at the rear end. The energy produced consists of a combustion
Compression - 2.5: 1
chamber, a fuel pump, a gas turbine, and an air compressor.
Specific fuel consumption 1.2 (mid thrust)
Atmospheric air enters the diffuser with the same velocity as
Compressor type - single stage centrifugal (radial outflow)
the propulsion unit. The function of a diffuser is to convert
Turbine type - single stage axial flow
the kinetic energy of entering air into its pressure energy. The
Ignition system - air gap, high voltage capacitor discharge
high-pressure air now enters the compressor, where its pressure
into igniter plug
is increased, and the temperature rises accordingly [2]. The
Lubrication system - fully recirculating auxiliary pump
high-pressure and temperature compressed air now enters the
Fuel system - metered return, spill type system, auxiliary
spherical combustion chamber in which the liquid fuel is
pump
injected through a ring through a pump. In the combustion
Fuel used - kerosene
chamber, the fuel burns at constant pressure by encountering
hot compressed air. The hot gases thus generated in the
combustion chamber now enter a gas turbine which expands to
produce power. It is with this power that the compressor and Procedure
the fuel pump are operated. Now, the exhaust gases from the 1) Assure 140 PSI of the compressed air supply.
turbine pass into the atmosphere through the expansion nozzle. 2) Verify the test rig’s electric supply, fuel level, and oil
As the exhaust gases pass through the nozzle, their pressure level.
energy is converted into kinetic energy. The gases leave the 3) Check The display after turning the key in the ignition.
nozzle with a very high velocity creating a reactive force or 4) Open the software and connect the data acquisition
thrust in the opposite direction. This moves the propellant unit system to the test rig.
forward. The turbine compressor compresses air in accordance 5) To start the engine, press the starter button.
with the Brayton cycle’s fundamental principles. Compression 6) Move the throttle lever to the proper setting for running
is depicted in steps 0–3, constant pressure heat addition in under typical operating conditions once the engine has
steps 3–4, and expansion through the turbine and nozzle in reached a speed above 14000 rpm. all pertinent infor-
steps 8–4. The engine’s net thrust is produced by the change mation and total the readings at a certain speed.
in kinetic energy of the gases between the exhaust and the 7) Repeat the process as many times as you can at the
inlet [3]. highest pace.
8) To slow the engine down, pull back on the throttle lever.
II. E XPERIMENTAL S ETUP
To turn the engine off, press the switch-off button.
The photograph of the test set-up is shown in Figure with 9) Switch off the ignition key and shut the compressed air
the different components marked. supply.
The test rig is equipped with an SR-30 engine that has a
sound suppressor at the inlet. The throttle valve is included III. T HEORY
with the testing apparatus. To gather the data, the entire test
rig is connected to a data acquisition system. The thrust of a gas turbine engine(here, turbojet engine)
can be divided into two parts:
Engine and accessory specifications - SR30 turbojet 1)Momentum Thrust
Engine diameter - 6.8 inches (17 cm) 2)Pressure Thrust
Engine length - 10.8 inches (27 cm) Actually, these two components are a net result of the
Design maximum thrust- 40 lbf (178 n) unbalance of surface forces acting across the engine.The
Mass flow rate - 1.1 lbs/s (0.5 kg/s) predominant surface forces are the pressure forces.Net Thrust
is given by: • Volume flow rate of air into the compressor,
ṁ = ρAVe kg/s
T = ṁa × (Ve − Vi ) + Ae × (Pe − Pa )
• Thrust due to outflow of air,
In this experiment exit pressure equals the ambient
pressure(Pe =Pa ).Therefore pressure thrust is neglected in this F2 = ṁVe kgm/s−2
experiment.Thus the engine SR30 is operating at its optimum
thrust condition. • Mach number of the airflow at this section,
Thus to calculate the thrust due to intake air at the Ve
M=√
compressor inlet, γRTe
• Diameter of the engine intake bell section at pitot static Net thrust of the engine is given by,
tube location,Dc = 6.604 cm
• Diameter of the tacho generator housing,d = 1.905 cm F = F2 − F1
• Cross sectional area at pitot static tube location,
Thrust specific Fuel Consumption is given by,
π
A = (Dc2 − d2 ) W eightof f uelburnedperhour
4 T.S.F.C =
T hrustF orce
• Pitot static gauge pressure at compressor inlet side,p1 =
p1,gauge IV. R ESULTS AND D ISCUSSION
• Pitot static absolute pressure at the compressor inlet side,
Observation values taken from the experiment are shown
pabs,1 = p1,gauge + patm in Table.I along with the calculated values. For analysing
the performance of the engine two performance parameters
• Temperature at the compressor inlet is T1
namely ’Thrust’ and ’TSFC’ are studied.
• Density of air corresponding to pabs,1 and T1 from air
Using the digital metres linked to the SR 30 Turbojet, the
table,is ρ
• Velocity of air at the compressor inlet,
r
2p1
Vi = m/s
ρ
• Volume flow rate of air into the compressor,
ṁ = ρAV kg/s
• Thrust due to intake air,
F1 = ṁVi kgm/s−2
• Mach number of the airflow at this section,
Vi
M=√
γRT
To find out the thrust due to exit gas at the nozzle exit,
• Diameter at nozzle exit,De = 5.6cm
• Cross-sectional area at the nozzle exit,
π Fig. 4: TSFC versus Thrust
A = (De2 )
4
data of the Inlet Compressor Temperature, Inlet Compressor
• Pitot static gauge pressure at nozzle exit side,pe =
Pressure, Pressure at the Nozzle exit, Exhaust Gas Tempera-
pexit,gauge
ture, and Rate of Fuel Consumption are determined for various
• Pitot static absolute pressure at the nozzle exit side,
rpm. It has been found that Thrust increases with increasing
pabs,e = p(exit,gauge) + patm rpm and TSFC drops with increasing rpm, which is consistent
with the literature of a turbojet engine. The graph for TSFC
• Temperature at Nozzle exit,is Te vs. thrust is then plotted in Fig.4. It has also been found that
• Density of exit gas corresponding to pabs,e and Te from thrust and TSFC are inversely proportional to one another. The
air table is ρ experiment is extremely helpful in determining the thrust that
• Velocity of air at the compressor inlet, is produced for a certain value of fuel consumption, which can
be helpful in determining the actual thrust that will develop
r
2pe
Vi = m/s when the engine is operated in actual conditions.
ρ
A similar variation is observed with TSFC versus mass flow can be seen from graph of TSFC vs Thrust that as the thrust
rate of fuel plot shown in Fig.5. A rectangular hyperbola was increases the TSFC value decreases and vice versa. Also it
observed showing that with increasing mass of fuel the thrust can be observed that as the mass flow rate increases the TSFC
decreases which is direct result of the non stichometric fuel value decreases. The plots generated are in accordance with
air mixture inside the combustion chamber. the literature of turbojet engine [3].
A PPENDIX
Sample Calculation-
To find out the thrust due to intake air at compressor inlet:
Diameter of the engine intake bell section Dc = 6.604cm
Diameter of the tacho generator housing, d = 1.905 cm
Cross sectional area
at intake,
A = π4 Dc2 − d2 = 31.403
Pitot static pressure at compressor inlet side,
p1 = 0.179P si
p1 = 102267P a
Temperature at compressor inlet, T1 = 34.290 C
Density of air corresponding to p1 and T1 from air table,
ρi = 1.16kg/m3
Velocity
q of air at compressor inlet,
Fig. 5: Fan speed versus Flow rate Vi = 2p ρi = 40.32m/s
1
Volume flow rate of air into the compressor,
Fig.6 shows variation of Thrust with the engine speed. A ṁi = ρi AVi = 0.146kg/s
linear variation could be seen between the two. It is noticed Thrust due to intake air, F1 = ṁi Vi = 5.886 N
that the trust increases to a point until the chocking condition Mach number of the air flow at this section,
is attained in the turbine stage, crossing this limit the thrust Vi
M1 = √γRT = 0.1147
1
decreases until the condition is reversed [4].
To find out the thrust due to exit gas at nozzle exit
Diameter at nozzle exit, De = 5.6 cm
Cross sectional area at nozzle exit, Ae = π4 De2 = 24.63cm2
Pitot static pressure at Nozzle exit,
pe = 0.583P si
pe = 105340P a
Temperature at Nozzle exit, Te = 529.300 C
Density of exit gas corresponding to pe and Te from air
table,
ρe = 0.457kg/m3
Velocity
q of air at Nozzle exit ,
Ve = 2p ρe = 132.63m/s
e
Volume flow rate of air at Nozzle exit,
ṁe = ρe Ae Ve = 0.15 m/s
Thrust due to outflow of air,F2 = ṁe Ve =19.89 N
Mach number of the air flow at this section,
Ve
M2 = √γRT = 0.23
Fig. 6: Thrust versus engine speed e
Net thrust of the engine
V. C ONCLUSION
F = F2 − F1 = 14.004N
The Thrust and TSFC is calculated with the help of readings
obtained from data acquisition system which is coupled to the
Thrust Specific Fuel Consumption
experimental setup. Graphs of TSFC vs mass flow rate, Thrust
W eight of the f uel burned per hour
vs RPM and TSFC vs Thrust were plotted using Matlab. It T.S.F.C = N et T hrust = 1.015Kg/hrN
S.No Compressor Compressor Turbine In- Turbine Nozzle Exit Fuel Con- Speed Thrust CompressorCompressorTurbine Turbine Exhaust
inlet exit let Pressure Exit Pressure sumption inlet Exit inlet Exit Gas
Pressure Pressure Pressure Temp
PSIG PSIG PSIG PSIG PSIG Gal/hr rpm Lbs ◦c ◦c ◦c ◦c ◦c
1 0.17 13.507 13.361 1.005 0.739 5.753 60486.203 10.991 34.16 160.26 620.22 615.65 530.7
2 0.141 11.196 11.062 0.865 0.6 5.095 56234.582 9.966 34.38 154.56 612.57 620.93 530.77
3 0.135 11.014 10.883 0.866 0.583 5.065 55873.25 9.93 34.29 157.62 610.29 619.34 529.3
4 0.113 8.849 8.651 0.73 0.436 3.971 51597.48 8.942 34.34 145.9 597.91 606.92 525.67
5 0.082 6.682 6.567 0.598 0.293 3.711 45490.711 8.137 34.45 135.42 587.49 590.39 516.42
6 0.073 5.47 5.365 0.512 0.207 3.338 41690.855 7.552 34.65 127.06 576.15 585.73 508.11
7 0.056 4.527 4.418 0.443 0.152 2.955 38364.5 7.186 34.65 121.36 557.21 576.21 502
TABLE I: Observation Table
R EFERENCES
[1] Performance Analysis on a Turbojet Engine. ndian Institute of Space
Science and Technolog.
[2] E. J. Manganiello, “Current nasa research in turbojet propulsion,” tech.
rep., 1970.
[3] R. D. Flack, Fundamentals of jet propulsion with applications, vol. 17.
Cambridge University Press, 2005.
[4] A. Smith, “Fluid mechanics, thermodynamics of turbomachinery. sl dixon.
pergamon press. 1966. 213 pp, diagrams. 25s.,” The Aeronautical Journal,
vol. 74, no. 719, pp. 915–915, 1977.
